JPH0653520B2 - Deployment of payload from shuttle using injection suppression device - Google Patents

Description

Description: BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a device for payload ejection such as launching a spacecraft or satellite from a launch vehicle, and in particular, the payload is gyroscopically launched from the launch vehicle. In particular, the present invention relates to a payload injection device including means for mechanically controlling an undesired force on a payload before injection.

Reference "Syncom IV Space Shuttle Orbital Fright Test Miss"
ion), publication number 66710V / December 1976, describes a device for launching a spacecraft from a launch vehicle. The device consists of a U-shaped or open-ended pedestal having an injection spring mechanism located on one side of the gantry, which mechanism includes a small projection protruding from one side of the spacecraft. Engage and eject the spacecraft. A pivot point is formed on the other side of the spacecraft, and this pivot point also has a projecting protrusion that sits on an engaging surface formed on the pedestal.

Although not described in this document, it is necessary that these spring mechanisms and pivots lie in a plane orthogonal to the spin axis of the spacecraft. Ideally, the center of mass of the spacecraft also lies in this plane. When the spring of the firing spring mechanism is unconstrained, tangential thrust acting in this plane is applied to the spacecraft. Assuming the above configuration to be accurate, when this spacecraft tangential thrust is applied, the spacecraft will rotate about this pivot, simultaneously producing nutation-free locomotion and rotational motion, Along with it, I go away from the gantry.

After this ejection is completed and the emission output disappears, the free-body motion of the spacecraft is a rolling motion that rolls up the virtual ramp, and maintains the applied linear momentum and angular momentum. The separating speed and rotation speed are
It is governed by inertial properties, diametrical dimensions, firing power, and injection stroke length.

United States Patent No. 43592 assigned to the assignee of this invention
The payload deployment system described in No. 01, "Payload Deployment with Linear and Angular Velocity from the Shuttle," is an improvement on the system described in the references cited above. The payload deployment system described in this patent employs a single point injection force application to provide a simple mechanization of payload deployment. The injection point and the reaction point between the spacecraft and the gantry surround the center of mass and the center of impact of the payload. This minimizes the need for elaborate restraint mechanisms at the point of action of force and the need to align the point of fire output exactly with the center of mass and the pivot point on which the reaction force acts.

This includes, at a minimum, a pair of pivot points that are located on one side of the spacecraft and are spaced apart from each other, and a firing force application point that is located on the other side of the spacecraft. To support. Each of these points forms a triangle surrounding the center of mass of the spacecraft. In such a three-point engagement system, the attitude of the spacecraft is determined when it is separated from the launch vehicle. Rotating the payload or spacecraft around fixed support points, such as pivots that are remote from each other, ensures more accurate payload deployment trajectories as well as physical separation during the ejection separation stage during ejection. I shall. Also, the effects of disturbances on the spacecraft attitude due to rocking of the liquid fuel while the spacecraft is being ejected and separated and in physical contact with the shuttle or launch vehicle are avoided.

The payload deployment system described in this patent is
While providing significant improvements for the payload ejection system, the pivotal support must support the payload structure, which in turn requires support for a structure that is less flexible in this patented design. Thermal and gravitational strains in these rigid structures can cause a phenomenon known as "gapping" at this pivot.

In addition, due to this inflexible support structure, a “racking” phenomenon may occur due to a statically indeterminate force having a large potential applied to the payload when the payload is deployed. Using a rigid support structure
It acts so as to combine the problems of the above-mentioned gapping and racking phenomena, and the "pivot bouncing phenomenon" is likely to occur during payload deployment. These phenomena are
It causes undesired rotation which will affect the attitude of the payload.

In a typical payload deployment, when a firing force is applied to the spacecraft by a spring-biased firing force actuator, a plurality of unlocking mechanisms that secure the spacecraft to the cradle exert forces to cancel it. The appearance of internal forces in these unlocking mechanisms also creates undesirable forces and moments that will be applied to the spacecraft. Therefore, eliminating these internal forces is beneficial for more accurate payload ejection.

SUMMARY OF THE INVENTION In order to improve upon the prior art payload or spacecraft deployment system, the present invention incorporates several configurations that define more accurate and regulated spacecraft ejection. The invention finds use in a device used to gyroscopically eject a spacecraft from a support structure. The support structure comprises a spacecraft support cradle having an open end, the opening being arranged to straddle the spacecraft.

The spacecraft has at least one mounting member on one side and at least two mounting members on the other side at longitudinally spaced positions. The mounting member is releasably engaged with the cradle and supports the spacecraft inside the cradle opening. The attachment members are generally spaced apart and form the approximate apex of a triangle surrounding the center of mass of the spacecraft. The apparatus includes an injection mechanism having a component disposed between the cradle and a mounting member for applying a tangential thrust to one side of the spacecraft at the position of the mounting member.

The invention comprises a separable injection suppression mechanism having one end mounted on a cradle and the other end mounted on a spacecraft. This ejection suppression mechanism suppresses the thrust force applied to the spacecraft for a predetermined time in order to eliminate the statically indeterminate force acting on the spacecraft prior to its ejection. The injection suppressing mechanism includes a separable elastic portion and a releasable clamp portion coupled to the elastic portion. The elastic portion comprises a spring biased anchor rod device and the releasable clamp portion comprises a spring biased split ball clamp device.

The present invention also contemplates the use of a flexible carrier coupled between the spacecraft and the cradle. The carrier includes one mounting member on one side and two mounting members on the other side. The carrier comprises a rigid portion containing one mounting member and a plurality of restraint mechanisms, and a flexible portion containing two mounting members. This carrier is used as a contact between the spacecraft and the cradle,
The ejection suppression mechanism contributes to the elimination of unwanted forces acting on the spacecraft prior to its ejection.

In general, the injection suppression mechanism acts as a tether that connects the gantry and the spacecraft cargo, creating a force-stabilizing state for this cargo, and creating a gap phenomenon at the pivot point prior to cargo deployment. Acts to remove the. The carrier is independent of the ejection structure, which is functionally easy to handle. With proper placement of the injection restraint mechanism, the pivot, the center of mass of the cargo and the ejection force actuator, the reaction force of the pivot and spring does not change during the release of the injection restraint, which results in bouncing of the pivot and cargo freight. Minimize deviation between the angular momentum vector and the static pivot axis of the cargo. Further, by using a single tether as a means of releasing a single support rather than releasing multiple support mechanisms during the ejection of cargo, the deformation energy of the structure allows these devices to move. Minimizes the chances that a cargo will experience uncontrolled momentum.

Description of the Preferred Embodiment Referring to FIG. 1, an injection suppression mechanism 30 according to the present invention.
Is shown as assembled to the spacecraft support structure 20. The support structure 20 comprises a spacecraft support pedestal 21 having an open end. The spacecraft 23 is arranged so that the open end of the support frame 21 straddles it.

The spacecraft 23 has at least one mounting member 2 on one side.
4 with at least two mounting members 2 on the opposite side
5 and 26 are provided at positions distant from each other in the longitudinal direction. The mounting members 24, 25, and 26 are detachably coupled to the support pedestal 21, and support the spacecraft 23 at the open end of the support pedestal 21. The attachment members 24, 25, 26 form a substantial vertex of a triangle surrounding the center of mass of the spacecraft 23 at a sufficient distance. The support structure 20 also includes an injection mechanism 27 having a portion located between the support platform 21 and the one mounting member 24. The ejection mechanism 27 is for applying a tangential thrust to the spacecraft 23 at the position of the mounting member 24.

For a better understanding of the injection structure within which this invention is employed, US Pat. No. 4,359,201 "Payload Deployment with Linear and Angular Velocity from the Shuttle"
(Payload Deployment from
Shuttle with Linear and
See Angular Velocity).
In addition, it is described as a reference also in this specification.

The injection suppressing mechanism 30 of the present invention is a separable device,
It has one end attached to the gantry 21 and the other end attached to the spacecraft 23. The ejection suppressing mechanism 30 suppresses the thrust force applied to the spacecraft 23 in order to cancel the indeterminate force acting on the spacecraft 23 prior to the ejection of the spacecraft 23.

As described in detail below, the injection restraint mechanism 30 generally comprises a separable elastic portion and a releasable clamp portion coupled thereto. The elastic portion may be a spring-biased anchor rod device and the releasable clamp portion may be a spring-biased split ball clamp device.

Referring to FIGS. 2a, 2b and 2c, there is shown a sequential description of the operation of the injection suppression mechanism 30 of FIG. In FIG. 2a, the spacecraft 23 is fixed to the fixed position of the pedestal 21 by a plurality of gantry fixing release mechanisms 31a, 31b, 31c. The ejection suppression mechanism 30 is in a compressed state because the spacecraft 23 is in a fixed state.

FIG. 2b shows that the gantry fixing release mechanism 31 is released, whereby a predetermined thrust force is applied to the cosmic ray 23 from the injection mechanism 27. In response to the thrust applied by the ejection mechanism 27, the spacecraft 23 moves slightly while rotating, and as a result, a downward force acts on the gantry 21 at the pivot points located at the attachment members 25 and 26.

The ejection suppression mechanism 30 suppresses the ejection stroke of the spacecraft 23 for a certain period of time, whereby the spacecraft 23 is completely engaged with the pivot point and "gapping" does not occur. During this time, indeterminate forces due to pivotal bouncing, thermal deformation, gravitational deformation, etc. of the support structure 20 are eliminated. In this method, due to the support structure that is difficult to bend due to being mounted on the mount, the spacecraft
The high static, static force applied to is removed.

In FIG. 2c, the spacecraft 23 is ejected from the support structure 20 and the ejection stroke is completed. In order to eject the spacecraft 20 completely, the ejection suppression mechanism 30 is separated and its elastic part remains attached to the spacecraft 23,
The clamp part remains attached to the pedestal 21.

To explain this in detail, the plurality of gantry fixing release mechanisms are released, and thereby the spring-biased injection mechanism raises and rotates the mounted object, whereby the pivot bearing portion is moved to the pad at the relevant position of the gantry 21. Press against. This operation continues until the injection suppression mechanism 30 reaches its balance limit. After a lapse of a predetermined time, the injection suppression mechanism 30 is separated, the spring force from the injection mechanism 27 is transmitted to the cargo, and a predetermined injection operation is performed.

Referring now to FIG. 3, one embodiment of the injection suppression mechanism 30 according to the present invention is illustrated. In this embodiment, the injection suppression mechanism 30 includes an elastic part 40, which comprises a cylindrical housing 41 in which a spring-biased anchor rod 42 is arranged. The anchor rod 42 is composed of a compression spring 43 and is connected to the adjusting rod 42. A base plate 45 for mounting the injection suppressing mechanism 30 on a spacecraft is provided at one end of the elastic portion 40. The end portion of the anchor rod 42 on the base plate side is connected to the base plate 45 by a pin 46, and this pin penetrates the end portion of the anchor rod 42 and the base plate 45. The other end of the anchor rod 42 has a ball shape.

The injection suppressing mechanism 30 also includes a clamp portion 50. The clamp portion 50 includes a spring-loaded split type ball clamp device, and has a housing 54 for attachment to the pedestal 21. The clamp portion 50 includes a split type ball clamp 51 for gripping the ball end portion of the anchor rod 42. FIG. 3 shows a state in which the split type ball clamp is split. The spring 52 is used in a compressed state and surrounds a separable rod 53. In FIG. 3, the rod 53 is shown in a separated state. The severable rod 53 can be a pyrotechnical connection or other severable mechanism such that the detachable ball clamp 51 can be actuated at any time.

More specifically, in other words, the injection suppressing mechanism 30 and, in other words, the tether are separable from the anchor rod 42 and the spring clamp device. The anchor rod 42 is attached to the spacecraft via a height adjusting device, and the height adjusting device can adjust the height to the contact point of the ball clamping device. The other end of the rod 42 is engaged with a split type ball clamp 51 attached to the pedestal 21. The split type ball clamp 51 is spring-biased so that the gripping of the anchor rod 42 is released by the separation of the separation rod 53.

The operation of the injection suppression mechanism 30 of FIG. 3 will be described subsequently. The anchor rod 42 is attached to the spacecraft by the base plate 45, while the clamp portion 50 is attached to the mount 21. The spacecraft is installed in place, with the anchor rod 42 moving to the down position and the spring 43 uncompressed. At this time, the split type ball clamp grips the ball end portion of the anchor rod 42. When the spacecraft is ready for injection, the plurality of unlocking mechanisms that have fixed the spacecraft to the gantry 21 are released.

Along with this, the ejection mechanism 27 attempts to eject the spacecraft from the gantry 21, but the ejection suppression mechanism 30 suppresses this operation on the way. The elasticity of the anchor rod device allows the spacecraft to move slightly, which causes the mounting member to press against the associated pad at each pivot that mounts the spacecraft to the cradle. In addition, cosmic rays are hydrostatically balanced, which eliminates the adverse effects of unwanted forces acting on the cosmic rays. After a lapse of a predetermined time, the clamp portion 50 is separated by a pyrotechnical operation or an operation similar thereto, whereby the grip of the anchor rod 42 is released. After that, the injection mechanism 2
7 ejects the spacecraft from the gantry 21.

Generally, spacecraft support structures and the like are of high rigidity to support the weight of the spacecraft and at the same time assist the ejection function. However, in order for the injection suppressing mechanism according to the present invention to operate most effectively, it is desirable that the rigidity is somewhat lower than that of a general one. Accordingly, the present invention provides a carrier structure that cooperates with an ejection suppression mechanism to assist in spacecraft deployment.

Referring to FIG. 4, there is shown a plan view of the carrier according to the present invention. The carrier 60 is attached to a perigee thrust motor 61 of the spacecraft 23. The carrier 60 includes a central portion 63 having high rigidity and a flexible pivot portion 64. The pivot 64 is cantilevered so that this part is somewhat elastic, which allows the pivot 64 to cooperate with the injection control mechanism of FIG. 3 to bring the spacecraft 23 into balance. Contribute.

The carrier 60 is arranged between the spaceship 23 and the gantry 21,
Used to secure the spacecraft to the gantry 21 prior to launch. For this fixing, a plurality of fixing mechanisms 65a to 65d are used.
Is used and its arrangement is shown in FIG. These locking mechanisms are mostly known from the prior art and are described in more detail in the documents mentioned in the background of the invention of this specification. Further, the mounting members 24, 25, 26 described with reference to FIG. 1 are shown in their positions but not in detail.

In the specific implementation shown in FIG.
Is mechanically attached to the perigid thrust motor 61, and this motor 61 is ejected from the spacecraft 23 when it is no longer needed, and this carrier is also ejected together with this motor.
This significantly reduces the weight of the spacecraft 23 and enables more efficient operation of the spacecraft 23 in orbit.

In order to better understand the principles of the present invention, the basic principles of operation of a so-called "frispy" firing system are described below. Reference is made to U.S. Pat. No. 4,359,201, "Payload Deployment with Linear and Angular Velocity from the Shuttle," which is cited in the Background of the Invention section of this specification, which describes the detailed working principle of a Frisbee injection system. ing.

This frispy injection system is designed to give the payload a velocity for the required separation relative to the launch rocket and a small angular velocity about the payload azimuth axis (physical symmetry axis). The angular momentum initially given to the payload contributes to restraining the flight attitude from changing as the payload flies away from the launch rocket.

This is because the payload is moving with the above-mentioned stable rotational movement. Direct precession of the payload azimuth axis as a result of undesired tip-off speeds on the payload during the release action on injection is prevented. In addition, postural instability that occurs in subsequent payload maneuvers is reduced.

The Frisbee injection system has three basic elements. On one side of the payload are two projections, or pivots, of which there are two elements, which are seated on engaging pivot supports located on the launch vehicle. Three
The pawl elements are located diametrically opposite these pivots. The third element is composed of a firing output generator mounted on the launch vehicle, and the firing output generator presses a pivot support portion arranged on the launch vehicle. A helical spring can be considered as the radiation output generation unit. The force exerted on the payload causes the pivot to be pressed against its seat when the spring produces a blast output. Payload is these 2
It rotates around the axis connecting the two pivots, increasing the straight line and rotation speed. When the spring finishes its stroke length from compression, the payload rotates away from the launch vehicle.

This rotation should in principle occur along the axis of symmetry of the payload. It is difficult to ensure that the axial position of the spring is coplanar with the axial position of the payload center of gravity. Also, any deviations in these positions result in a velocity across the axis of symmetry (chip-off velocity). For this reason, the system uses two pivots with some tolerance on the position of the center of gravity of the payload, which is due to any deviation between the center of gravity and the spring between the two pivots. This is because the pitch moment can be canceled out. However, in order to make the chip-off speed at the time of injection to zero, this 3
Each element of the point-engaged Frisbee injection system must be properly positioned with respect to the center of gravity of the system.

Rigid-body dynamics analysis of the Frisbee injection is insightful as to the guidance of the placement of these pivotal mechanisms. The steady-state value, which does not cause the undesirable pitch-direction tip-off velocity that occurs when the fulcrum reaction force is ejected, can be obtained from the analysis during the injection separation, where the force acting on the payload is the force. Suppose that it is as impulse rather than. Furthermore, this idealized analysis assumes that an axisymmetric, rigid payload is ejected from an object with infinite mass and inertia, the pivot point has infinite stiffness, and the ejection spring has its force Is assumed to act instantaneously.

The rotational and linear motion vector equations are related by a third equation linking angular velocity and linear velocity. This leads to three equations for the three unknowns, the dynamic reaction forces at the pivots R ₁ and R ₂ and the angular velocity ω _S.

In the special case where the pitch tip-off speed of the payload is zero in these movements, that is, it shows that the injection separation is performed under the optimum condition, and the equation expressing that state is as follows. Is used to relate the impulse of the spring to the impulse of the spring force.

I ₃ ω _S + r ₁ R ₁ + r ₂ R ₂ = R _S r _S (1) −hcm · R ₁ + (d−hcm) R ₂ = R _S (hcm−h _S ) (2) mr _S ω _S = R ₁ + R ₂ + R _S (3) These equations are considered to represent the roll, pitch, and linear velocity, respectively. For simplicity, r ₁ = r ₂
= R _S = r. Furthermore, two useful parameters related to the total mass of the payload can be defined as follows: β ² = I ₃ / mr ² and γ = (1−β ² ) / (1 + β ² ) ... (4) Here, I ₃ is the spin moment of inertia of the payload. Here, equations (1) and (2) can be solved for the compressive force required to create the optimum condition where the tip off speed is zero, and R ₁ / R _S = γ− (γ−1) hcm / d + h _S / d (5) R ₂ / R _S = (γ + 1) hcm / d−h _S / d (6) From this equation, the sum of reaction force impulses is γ of spring force impulses. Equivalent to double, which is to be expected for payloads that have rotational as well as linear motion.

Equations (5) and (6) derive the compressive force (here a positive value) over a range of values of hcm representing the center of gravity of the launch vehicle. For large hcm values or negative hcm values, one of these two pivots must be in tension and this must be the roof structure covering the pivot. If that is not possible. If the value of hcm causes a pulling reaction on the axis, then the equation for this state is solved for the tip-off velocity induced in it and the slope of the associated momentum vector.

Therefore, from this analysis, the tilt of the pivot reaction force and momentum vector is a function of the position of the center of gravity. Upon further analysis of these equations, the center of gravity of the system should fall within a region within the triangle that is somewhat smaller than the triangle defined by the three pivot points. If this criterion is met, it is possible to generate a dynamic reaction force on the pivot for injection separation with zero inclination.

The preference for the pivotal arrangement of the three-point injection system to produce these reaction forces resulting in low tilt injection separation has been described above. The injection suppression mechanism according to the invention acts as a suppression or towline element in the first of the two phases of the injection operation, so that the initial small consumption of energy of the spring preloads on these pivots. Contribute to that. Among these, in the first stage of the injection operation,
The payload may be constrained by the ejection restraint mechanism, which may be caused by less flexibility at the pivot fulcrum interface, or by racking or twisting of the plane forming the interface connecting the payload to the carrier. Stabilizes no transition fluctuations. The actual launch is then performed in the second stage by the mechanism releasing the restraint and ejecting the payload from the launch vehicle.

The reaction force of these pivots is generated such that the sum of these is γ times the spring force of the injection suppressing mechanism, and the injection suppressing mechanism is arranged outside the spacecraft in the radial direction toward the ejection output actuator.

This is as follows: rF _P + r _T F _T = rF _S (7) In a steady state, F _T + F _P + F _S = (1 + γ) F _S (8) Formula (7) is the radius of the injection suppression mechanism. Solved for the position, r _T / r = (F _S −F _P ) / F _T = (1−γ) / (1 + γ) = β ²
(9) In the pitch direction, it is desirable that the ratio between the forward direction and the stern pivot reaction force of the spacecraft is not affected by the restraint of the injection suppression mechanism, which means that the center of gravity about this mechanism is the axis of gravity. It becomes possible by arranging in a position. As a result, the injection suppressing mechanism does not have a pitch moment around the center of gravity of the payload while suppressing the payload from being ejected, and the pivot reaction force has the same value regardless of whether the injection suppressing mechanism is operating.

Thus, an injection suppression mechanism and a flexible carrier system for use in launching a spacecraft or other device from a launch vehicle have been described above. It is to be understood that the embodiments described herein merely illustrate some of the many specific embodiments that may find application in this invention. Obviously, numerous and numerous other devices can be readily devised by those skilled in the art within the scope of this invention.

BRIEF DESCRIPTION OF THE DRAWINGS Various features and advantages of the present invention will be more readily understood with reference to the following detailed description in connection with the accompanying drawings,
In the drawings, the same reference numbers represent the same structural elements,
FIG. 1 is a schematic view of an injection suppressing mechanism according to the present invention incorporated in a spacecraft injection device; FIGS. 2a, 2b and 2c show the operation of the injection suppressing device shown in FIG. FIG. 3 is a view showing an embodiment of an injection suppressing mechanism according to the present invention; FIG. 4 is a plan view of a carrier according to the present invention.

Front Page Continuation (72) Inventor Marhui, Zion, Earl, California, California 92708, Fountain Barry, Mount Dot 16600 (56) References US Patent 4359201 (US, A) US Patent 3380687 (US, A)

Claims (10)

[Claims] 1. An apparatus used for gyroscopically ejecting a spacecraft from a support structure, wherein the support structure comprises a spacecraft support mount having an open end, the spacecraft straddling the open end, and further comprising: The apparatus comprises an injection mechanism for applying a tangential thrust to one side of the spacecraft; has one end attached to the mount and the other end attached to the spacecraft, and prior to its ejection The apparatus comprising a separable injection suppression mechanism that suppresses applying the thrust to the spacecraft for a predetermined time in order to remove an indeterminate force acting on the spacecraft. 2. The apparatus according to claim 1, wherein the injection suppressing mechanism comprises a separable elastic portion and a releasable clamp portion coupled thereto. 3. An apparatus according to claim 2, wherein said elastic portion comprises a spring-biased strut device and said releasable clamp portion comprises a spring-biased split-type ball clamp device. 4. The spacecraft further comprises a carrier portion arranged in contact with the gantry, the carrier portion being located on one side of the one mounting member and the other one being located apart from the mounting member in the longitudinal direction. An apparatus according to claim 1, comprising said two mounting members on the side. 5. The apparatus according to claim 4, wherein the carrier portion includes a highly rigid portion composed of the one mounting member and a flexible portion composed of the two mounting members. 6. An apparatus for gyroscopically ejecting a spacecraft from a support structure, said support structure comprising a spacecraft support cradle having an open end, said spacecraft having at least one mounting member on one side thereof. And at least two attachment members on the other side located longitudinally apart, the spacecraft straddling the open end and the attachment members being releasably engageable with the gantry. The spacecraft is supported at an open end, the mounting member forms a triangular apex that sufficiently surrounds the center of mass of the spacecraft, and the apparatus further includes the one between the cradle and the one mounting member. An injection mechanism having a portion for applying a tangential thrust to one side of the spacecraft at the position of each mounting member; one end attached to the mount and the other end attached to the spacecraft Have This improved device comprises a separable injection suppression mechanism that suppresses applying the thrust to the spacecraft for a predetermined time in order to remove the statically indeterminate force acting on the spacecraft prior to the injection of the spacecraft. 7. The apparatus according to claim 6, wherein the injection suppressing mechanism comprises a separable elastic portion and a releasable clamp portion coupled thereto. 8. The apparatus according to claim 7, wherein said elastic portion comprises a spring-biased strut device and said releasable clamp portion comprises a spring-biased split-type ball clamp device. 9. The spacecraft further comprises a carrier portion arranged in contact with the gantry, the carrier portion being located on one side of the one mounting member and separated from the other mounting member in the longitudinal direction. 7. A device according to claim 6 comprising said two mounting members on the side. 10. The apparatus according to claim 9, wherein the carrier portion includes a highly rigid portion formed of the one mounting member and a flexible portion formed of the two mounting members.